NASA/AMES A-01 AIRFOIL (ames01-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: NASA/AMES A-01 AIRFOIL (ames01-il) Reynolds number: 100,000 Max Cl/Cd: 35 at α=10° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-ames01-il-100000.txt Download as CSV file: xf-ames01-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: NASA/AMES A-01 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.4401 0.09033 0.08590 -0.0089 1.0000 0.1356
-8.750 -0.4731 0.08491 0.08056 -0.0152 1.0000 0.1405
-8.500 -0.5197 0.07852 0.07419 -0.0231 1.0000 0.1412
-8.250 -0.5787 0.08546 0.08094 -0.0081 1.0000 0.1343
-8.000 -0.6169 0.07953 0.07483 -0.0186 1.0000 0.1409
-7.750 -0.5968 0.07514 0.07057 -0.0163 1.0000 0.1451
-7.500 -0.5930 0.07166 0.06702 -0.0177 1.0000 0.1538
-7.250 -0.5892 0.06735 0.06268 -0.0185 1.0000 0.1610
-7.000 -0.5916 0.06355 0.05869 -0.0201 1.0000 0.1738
-6.750 -0.5874 0.06075 0.05574 -0.0200 1.0000 0.1881
-6.250 -0.5672 0.05549 0.05057 -0.0166 1.0000 0.2196
-6.000 -0.5594 0.05331 0.04844 -0.0139 1.0000 0.2362
-5.750 -0.5580 0.04005 0.03297 -0.0146 1.0000 0.1131
-5.500 -0.5421 0.03594 0.02818 -0.0117 1.0000 0.0950
-5.250 -0.5269 0.03325 0.02522 -0.0097 1.0000 0.0937
-5.000 -0.5101 0.03126 0.02285 -0.0077 1.0000 0.0944
-4.750 -0.4916 0.02927 0.02051 -0.0059 1.0000 0.0945
-4.500 -0.4714 0.02748 0.01840 -0.0044 1.0000 0.0947
-4.250 -0.4507 0.02650 0.01702 -0.0027 1.0000 0.0967
-4.000 -0.4288 0.02439 0.01492 -0.0020 1.0000 0.1004
-3.750 -0.4059 0.02321 0.01365 -0.0010 1.0000 0.1032
-3.500 -0.3825 0.02222 0.01253 -0.0001 1.0000 0.1069
-3.250 -0.3591 0.02140 0.01158 0.0009 1.0000 0.1125
-3.000 -0.3355 0.02040 0.01070 0.0015 1.0000 0.1186
-2.750 -0.3117 0.01977 0.01003 0.0023 1.0000 0.1253
-2.500 -0.2889 0.01898 0.00940 0.0029 1.0000 0.1357
-2.250 -0.2663 0.01834 0.00886 0.0037 1.0000 0.1475
-2.000 -0.2435 0.01775 0.00839 0.0044 1.0000 0.1679
-1.750 -0.2207 0.01516 0.00801 0.0048 1.0000 0.5620
-1.500 -0.0949 0.01537 0.00904 -0.0096 1.0000 1.0000
-1.250 -0.0898 0.01539 0.00892 -0.0065 1.0000 1.0000
-1.000 -0.0814 0.01547 0.00887 -0.0037 1.0000 1.0000
-0.750 -0.0427 0.01565 0.00889 -0.0067 0.9945 1.0000
-0.500 0.0250 0.01582 0.00889 -0.0149 0.9820 1.0000
-0.250 0.0873 0.01592 0.00888 -0.0220 0.9720 1.0000
0.000 0.1490 0.01591 0.00880 -0.0287 0.9610 1.0000
0.250 0.2083 0.01582 0.00868 -0.0348 0.9487 1.0000
0.500 0.2587 0.01569 0.00854 -0.0387 0.9341 1.0000
0.750 0.2967 0.01556 0.00840 -0.0399 0.9168 1.0000
1.000 0.3207 0.01545 0.00830 -0.0383 0.8960 1.0000
1.250 0.3425 0.01524 0.00808 -0.0358 0.8756 1.0000
1.500 0.3603 0.01497 0.00781 -0.0325 0.8525 1.0000
1.750 0.3771 0.01455 0.00738 -0.0287 0.8278 1.0000
2.000 0.3935 0.01411 0.00692 -0.0249 0.7943 1.0000
2.250 0.4102 0.01365 0.00639 -0.0211 0.7422 1.0000
2.500 0.4250 0.01331 0.00564 -0.0163 0.6374 1.0000
2.750 0.4417 0.01396 0.00537 -0.0131 0.5070 1.0000
3.000 0.4637 0.01472 0.00560 -0.0119 0.4499 1.0000
3.250 0.4873 0.01534 0.00592 -0.0111 0.4199 1.0000
3.500 0.5114 0.01593 0.00628 -0.0104 0.3993 1.0000
3.750 0.5360 0.01648 0.00668 -0.0098 0.3827 1.0000
4.000 0.5607 0.01707 0.00712 -0.0093 0.3698 1.0000
4.250 0.5859 0.01762 0.00762 -0.0088 0.3580 1.0000
4.500 0.6111 0.01819 0.00817 -0.0083 0.3473 1.0000
4.750 0.6362 0.01885 0.00872 -0.0079 0.3380 1.0000
5.000 0.6616 0.01944 0.00936 -0.0075 0.3290 1.0000
5.250 0.6871 0.02015 0.01005 -0.0071 0.3209 1.0000
5.500 0.7124 0.02080 0.01072 -0.0067 0.3124 1.0000
5.750 0.7378 0.02157 0.01155 -0.0064 0.3049 1.0000
6.000 0.7631 0.02227 0.01230 -0.0061 0.2970 1.0000
6.250 0.7882 0.02310 0.01321 -0.0057 0.2895 1.0000
6.500 0.8132 0.02384 0.01401 -0.0054 0.2815 1.0000
6.750 0.8379 0.02473 0.01504 -0.0050 0.2739 1.0000
7.000 0.8627 0.02549 0.01585 -0.0047 0.2657 1.0000
7.250 0.8865 0.02643 0.01698 -0.0043 0.2574 1.0000
7.500 0.9112 0.02718 0.01772 -0.0039 0.2491 1.0000
7.750 0.9339 0.02808 0.01888 -0.0033 0.2395 1.0000
8.000 0.9579 0.02887 0.01964 -0.0029 0.2308 1.0000
8.250 0.9804 0.02950 0.02045 -0.0023 0.2203 1.0000
8.500 1.0020 0.03036 0.02152 -0.0016 0.2099 1.0000
8.750 1.0251 0.03083 0.02196 -0.0011 0.1997 1.0000
9.000 1.0474 0.03106 0.02223 -0.0004 0.1882 1.0000
9.250 1.0672 0.03138 0.02274 0.0006 0.1755 1.0000
9.500 1.0868 0.03164 0.02314 0.0016 0.1627 1.0000
9.750 1.1057 0.03182 0.02339 0.0028 0.1496 1.0000
10.000 1.1233 0.03209 0.02372 0.0040 0.1363 1.0000
10.250 1.1394 0.03269 0.02437 0.0054 0.1234 1.0000
10.500 1.1540 0.03366 0.02540 0.0068 0.1116 1.0000
10.750 1.1675 0.03485 0.02662 0.0083 0.1016 1.0000
11.000 1.1831 0.03620 0.02778 0.0095 0.0933 1.0000
11.250 1.1909 0.03821 0.03023 0.0115 0.0867 1.0000
11.500 1.2052 0.04002 0.03186 0.0126 0.0809 1.0000
11.750 1.2077 0.04234 0.03463 0.0149 0.0768 1.0000
12.000 1.2146 0.04444 0.03687 0.0166 0.0732 1.0000
12.250 1.2290 0.04736 0.03968 0.0173 0.0699 1.0000
12.500 1.2194 0.05021 0.04294 0.0202 0.0687 1.0000
12.750 1.2049 0.05334 0.04642 0.0229 0.0677 1.0000
13.000 1.1876 0.05701 0.05041 0.0247 0.0669 1.0000
13.250 1.1668 0.06136 0.05507 0.0254 0.0664 1.0000
13.500 1.1410 0.06675 0.06074 0.0246 0.0664 1.0000
13.750 1.1086 0.07378 0.06804 0.0219 0.0669 1.0000
14.000 1.0703 0.08298 0.07747 0.0165 0.0680 1.0000
14.250 1.0283 0.09451 0.08916 0.0091 0.0692 1.0000
14.500 0.8172 0.15803 0.15248 -0.0285 0.0927 1.0000
14.750 0.8212 0.16329 0.15775 -0.0298 0.0931 1.0000
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Polar data table (+)
Polar graphs
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